Blade outer air seal

ABSTRACT

A blade outer air seal for a gas turbine engine includes a first side surface, a second side surface, and a wall. The wall extends between the first side surface and the second side surface and has one or more holes formed therein. The holes are spaced from the first side surface and/or the second side surface and have areas between about 0.005% and 0.450% of a surface area of the blade outer air seal.

BACKGROUND

The invention relates to gas turbine engines, and more particularly toblade outer air seals (BOAS) for gas turbine engines.

Gas turbine engines operate according to a continuous-flow, Braytoncycle. A compressor section pressurizes an ambient air stream, fuel isadded and the mixture is burned in a central combustor section. Thecombustion products expand through a turbine section where bladed rotorsconvert thermal energy from the combustion products into mechanicalenergy for rotating one or more centrally mounted shafts. The shafts, inturn, drive the forward compressor section, thus continuing the cycle.Gas turbine engines are compact and powerful power plants, making themsuitable for powering aircraft, heavy equipment, ships and electricalpower generators. In power generating applications, the combustionproducts can also drive a separate power turbine attached to anelectrical generator.

The BOAS as well as turbine vanes are exposed to high-temperaturecombustion gases and must be cooled to extend their useful lives.Cooling air is typically taken from the flow of compressed air.Therefore, some of the energy that could be extracted from the flow ofcombustion gases must instead be expended to provide the compressed airused to cool the BOAS as well as the turbine vanes. Energy expended oncompressing air used for cooling the BOAS and turbine vanes is notavailable to produce power. Improvements in the efficient use ofcompressed air for cooling the BOAS and turbine vanes and/or materialsthat can better withstand the turbine operating environment can improvethe total efficiency of the turbine engine and extend the operationallife of the BOAS.

SUMMARY

A blade outer air seal for a gas turbine engine includes a first sidesurface, a second side surface, and a wall. The wall extends between thefirst side surface and the second side surface and has one or more holesformed therein. The holes are spaced from the first side surface and/orthe second side surface and have areas between about 0.005% and 0.450%of a surface area of the blade outer air seal.

A turbine section of a gas turbine engine includes an engine case, asupport connected to the engine case, and a blade outer air seal. Theblade outer air seal is mounted to the support and has a wall with abond coat and a thermal barrier coating. Both the bond coat and thethermal barrier coating have a radial thickness between 3% and 10% ofthe total radial thickness of the wall.

A blade outer air seal for a gas turbine engine includes a first sidesurface, a second side surface, a wall, and one or more forward hooks.The one or more forward hooks extend from the wall and at least one ofthe hooks has a slot therein that is offset relative to an axis ofsymmetry of the blade outer air seal.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partial cross-sectional view of an exemplary gas turbineengine.

FIG. 2 is an enlarged view of a turbine portion of the gas turbineengine shown in FIG. 1 with a BOAS mounted therein.

FIG. 3 is a perspective view of one embodiment of the BOAS.

FIG. 4 is a plane view of the outer radial surface of the BOAS of FIG.3.

FIG. 4A is a cross-sectional view of the BOAS of FIG. 4.

DETAILED DESCRIPTION

The present invention provides a BOAS design with higher convectiveefficiency and with improved durability due to improved corrosion andoxidation resistance. More particularly, the BOAS described hereinutilizes optimally sized holes in an outer diameter surface of a walland optimally sized passages within the wall to better control coolingair flow through the BOAS and thereby improve convective efficiency ofthe BOAS. These features improve the operational longevity of the BOAS.Additionally, the BOAS is adapted with features such as a non-symmetricslot and an angled hook wall that extends radially and axially to aid inassembly of the BOAS within a gas turbine engine.

An exemplary industrial gas turbine engine 10 is circumferentiallydisposed about a central, longitudinal axis or axial engine centerlineaxis 12 as illustrated in FIG. 1. The engine 10 includes in series orderfrom front to rear, low and high pressure compressor sections 16 and 18,a central combustor section 20 and high and low pressure turbinesections 22 and 24. In some examples, a free turbine section 26 isdisposed aft of the low pressure turbine 24. Although illustrated withreference to an industrial gas turbine engine, this application alsoextends to aero engines with a fan or gear driven fan, and engines withmore or fewer sections than illustrated.

As is well known in the art of gas turbines, incoming ambient air 30becomes pressurized air 32 in the compressors 16 and 18. Fuel mixes withthe pressurized air 32 in the combustor section 20, where it is burned.Once burned, combustion gases 34 expand through turbine sections 22, 24and power turbine 26. Turbine sections 22 and 24 drive high and lowrotor shafts 36 and 38 respectively, which rotate in response to thecombustion products and thus the attached compressor sections 18, 16.Free turbine section 26 may, for example, drive an electrical generator,pump, or gearbox (not shown).

It is understood that FIG. 1 provides a basic understanding and overviewof the various sections and the basic operation of an industrial gasturbine engine. It will become apparent to those skilled in the art thatthe present application is applicable to all types of gas turbineengines, including those with aerospace applications.

FIG. 2 is an enlarged view of a turbine section 22 and/or 24 of gasturbine engine 10 shown in FIG. 1 with a blade outer air seal (BOAS) 40disposed adjacent a turbine rotor blade 46. FIG. 2 illustrates BOAS 40,an engine case 42, stator vanes 44A and 44B, rotor blade 46, a BOASsupport 48, and a band segment 50. Vanes 44A and 44B include platforms43A and 43B.

BOAS 40 comprises an arcuate shroud segment with an inner diameter wallforming the outer diameter of the engine flow path through whichcombustion gases 34 pass. As will be discussed subsequently, passages(not numbered) extend through at least a portion of wall of BOAS 40 toprovide for cooling of BOAS 40 during operation. BOAS 40 is mountedwithin engine case 42 by forward and aft hooks, which engage BOASsupport 48 and vane platform 43B, respectively. BOAS support 48 and vaneplatforms 43A and 43B are in turn connected to engine case 42. Bandsegment 50 is positioned radially outward of BOAS 40 and extends betweenBOAS support 48 and vane platform 43B. Conformal seals such as w-sealsare disposed between vane platform 43B, BOAS support 48, and BOAS 40.

Rotor blade 46 comprises a single blade in a rotor stage disposeddownstream of combustor section 20 (FIG. 1). The rotor stage extends ina circumferential direction about engine centerline axis 12 (FIG. 1) andhas a plurality of rotor blades 46. During operation, combustion gases34 pass between adjacent rotor blades 46 and pass downstream to statorvane 44B. Rotor blade 46 is disposed radially inward of BOAS 40, withrespect to engine centerline axis 12 as shown in FIG. 1.

Stator vanes 44A and 44B are disposed axially forward and rearward ofBOAS 40, respectively. Like the rotor stage, the stator stages (of whichvanes 44A and 44B are a part) extend in a circumferential directionabout engine center line axis 12, and each stage has a plurality ofstator vanes. Vanes 44A and 44B include outer diameter platforms 43A and43B, respectively. Platforms 43A and 43B include features thatfacilitate the mounting stator vanes 44A and 44B to engine case 42.

In operation, the flow of combustion gases 34 impinges upon vanes 44Aand 44B and is aligned for a subsequent rotor stage. As the flow ofcombustion gases 34 passes through turbine blades 46 between a bladeplatform (not shown) and BOAS 40 the flow of combustion gases 40impinges upon rotor blade 46 causing the rotor stage to rotate aboutengine center line axis 12 (FIG. 1). BOAS 40 is mounted just radiallyoutward from rotor blade 46 tip and provides a seal against combustiongases 34 radially bypassing rotor blade 46. The flow of combustion gases34 exits rotor stage and enters stator vane stage passing vane 44B.

Engine case 42 and other components including vane platforms 43A and 43Bform plenums that can be used to communicate cooling air A to variouscomponents including BOAS 40, and in some embodiments, vanes 44A and44B. Generally, cooling air A is supplied to plenums from a source suchas high-pressure stage 18 and/or intermediate pressure stage ofcompressor (FIG. 1). Cooling air A passes through components such asBOAS 40 via passages (not shown) to provide for convection cooling.Thus, cooling air A provides desired cooling in order to increase theoperational life of BOAS 40.

FIG. 3 provides a perspective view of BOAS 40. BOAS 40 includes a wall51, an outer diameter surface 52, an inner diameter surface 54, a firstside surface 56, a second side surface 58, a forward surface 60, an aftsurface 62, a forward hooks 64, an aft hook 65, lateral hooks 66, holes68A and 68B, a slot 72, an angled wall 74 and a outer radial hooksurface 75.

Wall 51 of BOAS 40 has outer diameter surface 52, which extends betweenfirst side surface 56 and second side surface 58 and between forwardhooks 64 and aft hooks 65. Wall 51 has a total radial thickness Tbetween outer diameter surface 52 and inner diameter surface 54.Thickness T of wall 51 can vary from embodiment to embodiment, and caninclude a bond coat and/or a thermal barrier coating.

Inner diameter surface 54 is disposed on an opposing side of wall 51from outer diameter surface 52. Inner diameter surface 54 extendsbetween first side surface 56 and second side surface 58 and betweenforward surface 60 and aft surface 62. When BOAS 40 is installed in gasturbine engine 10 (FIG. 1), inner diameter surface 54 interfaces withand forms the outer diameter of the engine flow path through whichcombustion gases 34 pass (FIGS. 1 and 2). As will be discussedsubsequently, in some embodiments inner diameter surface 54 is formed byapplication of bond coat and thermal barrier coating.

First and second side surfaces 56 and 58 are disposed to either lateralside of BOAS 40. First and second side surfaces 56 and 58 intersect withforward surface 60. Forward surface 60 is disposed axially forward (withrespect to direction of flow of the combustion gases 34 through engineflow path) of forward hooks 64.

First and second side surfaces 56 and 58 also intersect with aft surface62, which extends radially inward of aft hook 65. Aft hook 65 extendsfrom wall 51 and is adapted to be received in a recess in vane platform43B (FIG. 2). Similarly, forward hooks 64 extend radially outward andforward from wall 51 and are adapted to be received in BOAS support 48(FIG. 2). Lateral hooks 66 extend radially outward from both wall 51 andforward hooks 64 over first side surface 56. Lateral hooks 66 overlapadjacent BOAS (not shown) when assembled.

Holes 68A and 68B (other holes not shown) are formed in wall 51 andextend through outer diameter surface 52 adjacent second side surfaces58. Holes 68A and 68B communicate with passages (FIG. 4A), which extendgenerally laterally through wall 51 from first side surface 56 to secondside surface 58. As will be discussed subsequently, holes, includingholes 68A and 68B, are sized to allow for the passage of optimal amountsof cooling air A (FIG. 2) into and through BOAS 40 in order to increasethe operational life of BOAS 40.

In the embodiment shown in FIG. 3, forward hooks 64 are separated byslots including slot 72. Slot 72 is offset relative to a lateral axis ofsymmetry A_(SM) of BOAS 40. Thus, BOAS 40, including forward hooks 64,has an asymmetric design in the lateral direction. Once assembled in gasturbine engine 10 (FIG. 1), slot 72 receives an anti-rotation feature(not shown) of BOAS support 48 (FIG. 2). Slot 72 and anti-rotationfeature prevent lateral movement (movement circumferentially aroundrotor stage within circumferential engine case 42) of BOAS 40.

Angled wall 74 extends radially and axially from outer radial hooksurface 75 to connect to outer diameter surface 52 of wall 51. Angledwall 74 provides for ease of identification of BOAS 40 during assemblyand disassembly processes.

FIG. 4 shows a plane view of the outer diameter of BOAS 40. FIG. 4Ashows a cross-sectional view of BOAS 40. As illustrated in FIGS. 4 and4A, BOAS 40 includes outer diameter surface 52, inner diameter surface54 (FIG. 4A), first side surface 56 (FIG. 4), second side surface 58(FIG. 4), forward surface 60, aft surface 62, forward hooks 64, aft hook65, lateral hooks 66 (FIG. 4), slot 72 (FIG. 4), passages 70 (FIGS. 3and 4A), angled wall 74 and outer radial hook surface 75. In addition toholes 68A and 68B, FIG. 4 illustrates holes 68C, 68D, 68E, and 68F. FIG.4A illustrates a bond coat 76 and a thermal barrier coating 78.

As illustrated in FIG. 4, holes 68A, 68B, and 68C extend through outerdiameter surface 52 adjacent second side surface 58 and holes 68D, 68E,and 68F extend through outer diameter surface 52 adjacent first sidesurface 56. As discussed, holes 68A, 68B 68C, 68D, 68E, and 68Fcommunicate with passages 70 (FIGS. 3 and 4A). Holes 68A, 68B 68C, 68D,68E, and 68F having varying diameters and are sized to allow for thepassage of optimal amounts of cooling air A (FIG. 2) into and throughBOAS 40 in order to increase the operational life of BOAS 40. Thus, inone embodiment each hole 68A, 68B 68C, 68D, 68E, and 68F has an areabetween about 0.005% and 0.45% of the surface area of BOAS 40 (asmeasured along a plane extending between first side surface 56, secondside surface 58, forward surface 60, and aft surface 62). In a furtherembodiment each hole 68A, 68B 68C, 68D, 68E, and 68F has an area betweenabout 0.020% and 0.30% of the surface area of BOAS 40 (as measured alonga plane extending between first side surface 56, second side surface 58,forward surface 60, and aft surface 62).

As shown in FIG. 4A, in one embodiment each passage 70 has a radialheight H₁ that comprises between about 30% to 40% of the total radialthickness T of wall 51. Passages 70 axial length L in total (between allsix passages) comprises between about 75% and 85% of the axial length ofBOAS 40 between forward surface 60 and aft surface 62.

FIG. 4A additionally shows bond coat 76, which is added to the wall 51.In one embodiment, bond coat 76 comprises a metallic coating thatprovides for increased oxidation and corrosion resistance. Bond coat 76can be a nickel alloy layer applied using plasma or high velocityoxy-fuel deposition processes. In one embodiment, bond coat 76 has aradial thickness H₂ between about 3% and 10% of the total radialthickness T of wall 51.

Thermal barrier coating 78 can be applied to bond coat 76 to form innerradial surface 52. In one embodiment, thermal barrier coating comprisesa ceramic layer that simultaneously provides thermal insulation andabradability and has a thickness H₃ between about 3% and 10% of thetotal radial thickness T of wall 51. The thermal bearing coating 78 canbe applied using plasma deposition or other known methods.

BOAS 40 can be constructed of metallic material such as a nickel basealloy that offers high temperature strength and hot corrosionresistance. In one embodiment, BOAS 40 is formed of a single crystalalloy that is cast and directionally solidified. The alloy canadditionally be heat treated at various temperature ranges for varyingdurations as desired.

The present invention provides a BOAS design with higher convectiveefficiency and with improved durability due to improved corrosion andoxidation resistance. More particularly, the BOAS described hereinutilizes optimally sized holes in an outer diameter surface of a walland optimally sized passages within the wall to better control coolingair flow through the BOAS and thereby improve convective efficiency ofthe BOAS. These features improve the operational longevity of the BOAS.Additionally, the BOAS is adapted with features such as a non-symmetricslot and an angled hook wall that extends radially and axially to aid inassembly of the BOAS within a gas turbine engine.

DISCUSSION OF POSSIBLE EMBODIMENTS

The following are non-exclusive descriptions of possible embodiments ofthe present invention.

A blade outer air seal for a gas turbine engine includes a first sidesurface, a second side surface, and a wall. The wall extends between thefirst side surface and the second side surface and has one or more holesformed therein. The holes are spaced from the first side surface and/orthe second side surface and have areas between about 0.005% and 0.450%of a surface area of the blade outer air seal.

The blade outer air seal of the preceding paragraph can optionallyinclude, additionally and/or alternatively, any one or more of thefollowing features, configurations and/or additional components.

The one or more holes have areas between about 0.02% and 0.30% of asurface area of the blade outer air seal.

The one or more holes comprise six holes with three holes positionedadjacent the first side surface and three holes positioned adjacent thesecond side surface.

Internal passages extend through the wall from the first side surface tothe second side surface, and wherein the one or more holes communicatewith the internal passages.

The six holes comprise one hole for each of the internal passages.

Each of the internal passages has a radial height between 30% to 40% ofan total radial thickness the wall.

The internal passages together have an axial length that comprisesbetween 75% and 85% of the axial length of the wall.

One or more forward hooks extend from the wall, and at least one of theforward hooks has a slot therein that is offset relative to an axis ofsymmetry of the blade outer air seal.

At least one of the forward hooks has an angled wall that extends froman outer radial surface of the at least one of the forward hooks to thewall.

The wall includes a bond coat, wherein the bond coat has a radialthickness between 3% and 10% of the total radial thickness of the wall.

The wall has a thermal barrier coating applied to an inner radialsurface thereof, wherein the thermal barrier coating has a radialthickness between 3% and 10% of the total radial thickness of the wall.

A blade outer air seal for a gas turbine engine includes a first sidesurface, a second side surface, and a wall. The wall extends between thefirst side surface and the second side surface and has a bond coat and athermal barrier coating. Both the bond coat and the thermal barriercoating have a radial thickness between 3% and 10% of the total radialthickness of the wall.

The blade outer air seal of the preceding paragraph can optionallyinclude, additionally and/or alternatively, any one or more of thefollowing features, configurations and/or additional components.

One or more holes are formed in the wall and are spaced from at leastone of the first side surface or second side surface, and the one ormore holes have areas between about 0.005% and 0.450% of a surface areaof the blade outer air seal;

One or more holes are formed in the wall and are spaced from at leastone of the first side surface or second side surface, and the one ormore holes have areas between about 0.020% and 0.30% of a surface areaof the blade outer air seal;

The one or more holes comprise six holes with three holes positionedadjacent the first side surface and three holes positioned adjacent thesecond side surface.

Internal passages that extend through the wall from the first sidesurface to the second side surface, and the one or more holescommunicate with the internal passages.

Each of the internal passages has a radial height between 30% to 40% ofan total radial thickness the wall.

The internal passages together have an axial length that comprisesbetween 75% and 85% of the axial length of the wall.

One or more forward hooks extend from the wall, wherein at least one ofthe forward hooks has a slot therein that is offset relative to an axisof symmetry of the blade outer air seal; and

At least one of the forward hooks has an angled wall extends from anouter radial surface of the at least one of the forward hooks to thewall.

A blade outer air seal for a gas turbine engine includes a first sidesurface, a second side surface, a wall, and one or more forward hooks.The one or more forward hooks extend from the wall and at least one ofthe hooks has a slot therein that is offset relative to an axis ofsymmetry of the blade outer air seal.

The blade outer air seal of the preceding paragraph can optionallyinclude, additionally and/or alternatively, any one or more of thefollowing features, configurations and/or additional components.

At least one of the forward hooks has an angled wall that extends froman outer radial surface of the at least one of the forward hooks to thewall.

While the invention has been described with reference to an exemplaryembodiment(s), it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment(s) disclosed, but that theinvention will include all embodiments falling within the scope of theappended claims.

1. A blade outer air seal for a gas turbine engine, the blade outer airseal comprising: a first side surface; a second side surface; and a wallextending between the first side surface and the second side surface,wherein one or more holes are formed in the wall and are spaced from atleast one of the first side surface or second side surface, and whereinthe one or more holes have areas between about 0.005% and 0.450% of asurface area of the blade outer air seal.
 2. The blade outer air seal ofclaim 1, wherein the one or more holes have areas between about 0.02%and 0.30% of a surface area of the blade outer air seal
 3. The bladeouter air seal of claim 1, wherein the one or more holes comprise sixholes with three holes positioned adjacent the first side surface andthree holes positioned adjacent the second side surface.
 4. The bladeouter air seal of claim 1, further comprising internal passages thatextend through the wall from the first side surface to the second sidesurface, and wherein the one or more holes communicate with internalpassages.
 5. The blade outer air seal of claim 4, wherein the one ormore holes comprise six holes with one hole for communicating with eachone of the internal passages
 6. The blade outer air seal of claim 4,wherein each of the internal passages has a radial height between 30% to40% of a total radial thickness the wall.
 7. The blade outer air seal ofclaim 5, wherein the internal passages together have an axial lengththat comprises between 75% and 85% of the axial length of the wall. 8.The blade outer air seal of claim 1, further comprising one or moreforward hooks that extend from the wall, wherein at least one of theforward hooks has a slot therein that is offset relative to an axis ofsymmetry of the blade outer air seal.
 9. The blade outer air seal ofclaim 8, wherein at least one of the forward hooks has an angled wallthat extends from an outer radial surface of the at least one of theforward hooks to the wall.
 10. The blade outer air seal of claim 1,wherein the wall has a thermal barrier coating applied to an innerradial surface thereof, wherein the thermal barrier coating has a radialthickness between 3% and 10% of a total radial thickness of the wall,wherein the wall includes a bond coat, and wherein the bond coat has aradial thickness between 3% and 10% of a total radial thickness of thewall.
 11. A blade outer air seal for a gas turbine engine, the bladeouter air seal comprising: a first side surface; a second side surface;and a wall extending between the first side surface and the second sidesurface, wherein the blade outer air seal has a wall with a bond coatand a thermal barrier coating, and wherein both the bond coat and thethermal barrier coating have a radial thickness between 3% and 10% ofthe total radial thickness of the wall.
 12. The blade outer air seal ofclaim 11, one or more holes are formed in the wall and are spaced fromat least one of the first side surface or second side surface, andwherein the one or more holes have areas between about 0.005% and 0.450%of a surface area of the blade outer air seal.
 13. The blade outer airseal of claim 12, wherein the one or more holes have areas between about0.02% and 0.30% of a surface area of the blade outer air seal.
 14. Theblade outer air seal of claim 13, wherein the one or more holes comprisesix holes with three holes positioned adjacent the first side surfaceand three holes positioned adjacent the second side surface.
 15. Theblade outer air seal of claim 11, further comprising internal passagesthat extend through the wall from the first side surface to the secondside surface, and wherein the one or more holes communicate withinternal passages.
 16. The blade outer air seal of claim 15, whereineach of the internal passages has a radial height between 30% to 40% ofa total radial thickness the wall.
 17. The blade outer air seal of claim15, wherein the internal passages together have an axial length thatcomprises between 75% and 85% of the axial length of the wall.
 18. Theblade outer air seal of claim 11, further comprising one or more forwardhooks that extend from the wall, wherein at least one of the forwardhooks has a slot therein that is offset relative to an axis of symmetryof the blade outer air seal, and wherein at least one of the forwardhooks has an angled wall that extends from an outer radial surface ofthe at least one of the forward hooks to the wall.
 19. A blade outer airseal for a gas turbine engine, the blade outer air seal comprising: afirst side surface; a second side surface; a wall extending between thefirst side surface and the second side surface; and one or more forwardhooks extending from the wall, wherein at least one of the forward hookshas a slot therein that is offset relative to an axis of symmetry of theblade outer air seal.
 20. The blade outer air seal of claim 19, whereinat least one of the forward hooks has an angled wall that extends froman outer radial surface of the at least one of the forward hooks to thewall.